Segmented ceramic matrix composite turbine airfoil component

ABSTRACT

A segmented component for use with a gas turbine engine comprises a radially extending gas path portion. The gas path portion is for interacting with gas flow from the gas turbine engine. The gas path portion comprises a forward portion forming a leading edge of a stationary vane, an aft portion forming a trailing edge of the stationary vane, and a plurality of middle portions forming a pressure side and a suction side of the stationary vane. The component is divided into axially aligned segments comprising a forward segment, an aft segment, and a plurality of middle segments disposed between the forward segment and the aft segment. The middle segments comprise radially elongate ceramic matrix composite material plates.

CROSS-REFERENCE TO RELATED APPLICATION(S)

This application is a divisional of Ser. No. 12/361,192, filed Jan. 28,2009, entitled “SEGMENTED CERAMIC MATRIX COMPOSITE TURBINE AIRFOILCOMPONENT”.

BACKGROUND

The present invention is directed to airfoil components for gas turbineengines and, more particularly, to ceramic matrix composite turbineblades, vanes and platforms.

Gas turbine engines comprise one or more rotating turbines that are usedto extract energy from a high velocity and high temperature gas flowproduced within the gas turbine engine. The turbines are comprised of aplurality of radially extending airfoil blades that are connected attheir inner diameter ends to a rotor, which is connected to a shaft thatrotates within the engine as the blades interact with the gas flow. Therotor typically comprises a disk having a plurality of axial retentionslots that receive mating root portions of the blades to prevent radialdislodgment. Blades typically also include integral inner diameterplatforms that prevent the high temperature gases from escaping throughthe radial retention slots. Between the turbine blades are disposed aplurality of radially extending stationary airfoil vanes, which aretypically supported by inner and outer diameter shrouds that aresuspended from an outer engine case and supported by an inner structure,respectively. During operation of the engine, the turbine blades andvanes are subjected to high heat from the gas. Additionally, the bladesare subjected to high stresses from rotational forces. It is, therefore,a constant design challenge to develop materials for turbine blades andvanes that are more heat resistant to reduce cooling demands, andlighter to increase propulsive efficiencies in aircraft engines.

Typically, turbine blades and vanes are fabricated from high strengthalloys as single pieces, with integral roots, platforms and shrouds.More recent turbine blade designs have attempted to incorporate ceramicmatrix composite (CMC) materials, which are lightweight, heat resistantand strong. CMC material comprises a ceramic fabric that is infused witha liquid ceramic matrix. The ceramic fabric is preformed to a desiredshape and the matrix solidifies within the fabric to produce a parthaving the lightweight and heat resistance characteristics of the matrixand the strength characteristics of the fabric. Production of thick CMCmaterial parts is constrained because of manufacturing limitations ininfusing the liquid matrix into deep layering of the preformed fabric.Inadequately infused liquid matrix produces porosity within thecomponent that limits the heat transfer capabilities within the matrixand provides an initiation point for crack propagation. Furthermore, dueto the two-dimensional nature of the ceramic fabric, difficulties arisein producing parts having complex three-dimensional shapes. For example,it is difficult to produce CMC material turbine blades having both aradially extending airfoil component and an axially extending platformcomponent.

It is, however, desirable to use CMC material despite thesecomplexities, as CMC materials weigh approximately one third of theweight of typical metal alloys used for turbine components, while havingmuch higher temperature limitations. As such, production methods havebeen developed that attempt to overcome the aforementioned manufacturingissues. However, previous attempts at producing CMC material componentshave resulted in complex designs that limit the benefits of using CMCmaterial in turbine blades. For example, one method of producing a vaneinvolves radially stacking numerous layers of CMC material in a radialdirection to obtain an airfoil shape. The stack is compressed withmechanical tensioning means to obtain the desired tensile strength andto prevent the layers from separating. Such a vane design is impracticalfor turbine blades because the stack extends radially in the directionin which severe stresses are generated within a rotating turbine blade.Furthermore, the mechanical tensioning means typically comprisesthreaded fasteners fabricated from an alloy that requires cooling, thuslimiting the temperature benefits of using CMC material in a hot sectionof a gas turbine engine. There is, therefore, a need for improved CMCmaterial turbine components and methods for fabricating the same.

SUMMARY

The present invention is directed to a segmented component for use witha gas turbine engine. The segmented component comprises a radiallyextending gas path portion. The gas path portion is for interacting withgas flow from the gas turbine engine. The gas path portion comprises aforward portion forming a leading edge of a stationary vane, an aftportion forming a trailing edge of the stationary vane, and a pluralityof middle portions forming a pressure side and a suction side of thestationary vane. The component is divided into axially aligned segmentscomprising a forward segment, an aft segment, and a plurality of middlesegments disposed between the forward segment and the aft segment. Themiddle segments comprise radially elongate ceramic matrix compositematerial plates.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 shows a cut-away perspective view of a rotor disk having radialretention slots connected with segmented ceramic matrix composite (CMC)turbine blade components of the present invention.

FIG. 2 shows an assembled view of segments from a removable CMC bladeairfoil of FIG. 1.

FIG. 3 shows an assembled view of segments from a removable CMC bladeplatform of FIG. 1.

FIG. 4 shows a partially exploded view of the removable CMC bladeplatform segments of FIG. 3.

FIG. 5A shows a CMC blade platform segment of FIG. 4 fabricated fromuni-directional ceramic fibers.

FIG. 5B shows a CMC blade platform segment of FIG. 4 fabricated frombi-directional fibers.

FIG. 6 shows a partially cut away view of an embodiment of a retentionmechanism for use with the removable segmented CMC blade platform ofFIGS. 3 and 4.

FIG. 7 shows another embodiment of a retention mechanism for use with aremovable CMC blade platform of the present invention.

FIG. 8 shows an embodiment of a stationary segmented CMC vane havinginner and outer diameter shrouds supported by a tongue and groove means.

FIG. 9 shows another embodiment of a stationary segmented CMC vanehaving inner and outer diameter shrouds supported by retention lugmeans.

FIG. 10 shows a cut away view of the stationary segmented CMC vane ofFIG. 9 showing a retention lug interacting with retention tabs at anouter diameter shroud.

DETAILED DESCRIPTION

FIG. 1 shows a cut-away perspective view of rotor 10 connected tosegmented ceramic matrix composite (CMC) material rotor blade components12 and 14 of the present invention. Rotor 10 comprises an annular body,such as a disk, that is configured to rotate about an axial centerlinewithin a gas turbine engine. Rotor 10 includes an inner diameterconfigured to be connected to a shaft in a gas turbine engine in anyconventional manner, and an outer diameter configured to connect tosegmented CMC blade 12 and segmented CMC platform 14 at retention slots16 and 18, respectively. Blade 12 and platform 14 each include a rootportion and a gas path portion. Specifically, blade 12 includes bladeroot 12A, which is connected to slot 16, and airfoil 12B. Likewise,platform 14 includes platform root 14A, which is connected to slot 18,and stage 14B. Blade 12 and platform 14 comprise one cluster of aplurality of clusters that would typically be disposed about the outerperiphery of rotor 10. As such, the plurality of blades comprises anannular array of airfoils that extend from rotor 10 in a radialdirection with respect to the engine centerline. The plurality ofplatforms comprises an annular ring that connects the plurality ofairfoils and extends circumferentially around the engine centerline. Inone embodiment of the invention, rotor 10 comprises a rotor disk for usein a hot section of the gas turbine engine such as in a low pressureturbine stage or a high pressure turbine stage. To enable use in hightemperature environments associated with hot sections of gas turbineengines, blade 12 and platform 14 are fabricated from CMC material. Aswill be laid out in greater detail below, the present inventionmaximizes the benefits of the CMC material by first separating blade 12and platform 14 into separate components, and then dividing eachcomponent into longitudinal segments. The CMC material is fullydensified to achieve optimal heat transfer characteristics, and isoriented to achieve optimal strength characteristics. As such, thelongitudinal segments of CMC material can be used in hot sections of agas turbine engine to form various components, such as blades,platforms, vanes and shrouds. FIGS. 1-7 discuss rotary turbinecomponents, such as blade 12 and platform 14. FIGS. 8 and 9 discussstationary turbine components, such as vanes and shrouds.

Blade 12 and platform 14 of the present invention are produced asseparate pieces to minimize angular geometries within each componentthat give rise to difficulties in producing CMC material components. Aplurality of blades 12 and platforms 14 are individually assembled torotor 10, after which each blade 12 is outfitted with an inner diametershroud. Blade root 12A of blade 12 and platform root 14A of platform 14are assembled with rotor 10 at slots 16 and 18, respectively. Platformslot 18 is positioned adjacent blade slot 16 such that an alternatingpattern of slots is formed in the periphery of rotor 10. Blade slot 16is generally larger than platform slot 18 to accommodate the greaterradial forces generated by the rotation of blade 12 than the forcesgenerated by platform 14 during rotation of rotor 10. Blade root 12A isinserted axially into slot 16 to radially restrain blade 12. Platformroot 14A is inserted axially into slot 18 to radially restrain platform14. In the embodiment shown, root 12A and slot 16 include fir treeengagement 20 in which ribs on root 12A engage channels on slot 16 thatpermit root 12A to be inserted into slot 16 axially, but that restrainradial movement of blade 12. In the embodiment shown, root 14A and slot18 include dovetail engagement 22 in which hooks on slot 18 engagesidewalls of root 14A that permit root 14A to be inserted into slot 18axially, but that restrain radial movement of platform 14. In otherembodiments, blade root 12A and platform root 14A may include otherradial retention configurations, such as dovetails or fir trees. Bladeroot 12A, platform root 14A, slot 16 and slot 18 are shaped to allow forthermal growth between roots 12A and 14A and slots 16 and 18 from heatgenerated during operation of the gas turbine.

Gas path portion 12B of blade 12 and gas path portion 14B of platform 14extend from root portions 12A and 12B, respectively, to form an airfoiland an inner diameter platform. Airfoil 12B extends radially from root12A and is shaped to include a pressure side and a suction side toextract energy to a flow of gases traveling perpendicular to blades 12.Stage 14B extends radially from root 14A and is shaped to provide a sealbetween a pressure side and a suction side of adjacent blades. Stage 14Bseals the base of airfoil 12B and prevents gas flowing over airfoil 14Bfrom entering slots 16 and 18. Stage 14B also prevents other fluids,such as cooling air circulated through rotor 10 from entering the flowof gas across airfoil 12B. The geometries and curvatures of the abuttingsurfaces of stage 14B and airfoil 12B are thus mating such that acontinuous inner diameter surface is formed near the transition regionbetween root portion 12A and airfoil portion 12B, similar to that ofconventional, integral platform blades. However, the present inventionsplits the platforms from the airfoils along the contour of the blade,rather than splitting an integrated platform between airfoils along astraight line. The specific geometries of stage 14B and airfoil 12B areselected based on aerodynamic needs, such as the shape of the pressureand suction sides, and radial retention need, such as the shape ofengagements 20 and 22. The geometries of stage 14B and airfoil 12B are,however, also selected so that they may be easily divided into thinsegments to facilitate fabrication from CMC material.

Blade 12 and platform 14 are further divided into thin, radiallyextending CMC material segments that better align ceramic fibers withinthe material and that better enable infusion of a ceramic liquid intothe fibers. Blade 12 is divided into segments comprising leading edgesegment 24A, trailing edge segment 24B and a plurality of middlesegments 24C. Platform 14 is divided into leading edge segment 26A,trailing edge segment 26B and a plurality of middle segments 26C. In theembodiment shown, blade 12 includes seven middle segments and platform14 includes ten middle segments. In other embodiments, however, othernumbers of middle segments can be used.

FIG. 2 shows an assembled view of segments 24A-24C from removable CMCblade 12 of FIG. 1. Each segment of blade 12 includes a section of root12A and airfoil 12B. Thus, each segment includes a gas path portionconfigured to impinge the hot gas of the gas turbine engine, and aradial retention portion configured to engage rotor 10. Root 12Acomprises a generally uniform cross sectional area in the axialdirection. As such, each segment of root 12A comprises a generallyequivalent root portion. When segments 24A, 24B and 24C are axiallyaligned, root 12A can be axially inserted into slot 16 so that fir treeengagement 20 is properly seated. Specifically, ribs 28 of root 12A areaxially seated within channels 30 of slot 16 to prevent displacement ofblade 12 in the radial direction.

Airfoil 12B includes leading edge 32, trailing edge 34, concave pressureside 36 and convex suction side 38, and extends in a generally arcuate,or airfoil shape, manner in the axial direction. Airfoil 12B alsoextends generally radially from root 12A such that each of segments 24A,24B and 24C include a small slice of the airfoil profile. Leading edgesegment 24A includes a rounded forward surface that forms leading edge32 and is designed to impinge an oncoming gas stream. Leading edgesegment 24A also includes a planar rear surface designed to abut one ofmiddle segments 24C. Trailing edge segment 24B includes a rounded rearsurface that forms trailing edge 34 and is designed to minimize flowseparation of the gas stream. Trailing edge segment 24B also includes aplanar forward surface designed to abut one of middle segments 24C.Middle segments 24C include planar forward and rear surfaces so thatthey can be axially stacked or aligned between leading edge segment 24Aand trailing edge segment 24B to complete the shape of an airfoil. Thesides of middle segments 24C include small segments of pressure side 36and suction side 38 to complete the airfoil shape. Additionally, eachsegment can be split into pressure side and suction side segment halvesto facilitate assembly with rotor 10 and platform 14, and to producehollow airfoils to minimize weight. For example, leading edge segment24A includes pressure side 40 and suction side 42, which comprise matingcurved segments that meet to form interior cavity 44. Interior cavity 44can be used to provide cooling air to blade 12 or to reduce the weightof blade 12.

FIG. 3 shows an assembled view of segments 26A-26C from removable CMCblade platform 14 of FIG. 1. FIG. 4, which is discussed concurrentlywith FIG. 3, shows a partially exploded view of segments 26A-26C fromremovable CMC blade platform 14 of FIG. 3. Each segment of platform 14includes a section of root 14A and stage 14B. Thus, each segmentincludes gas path portion configured to impinge the hot gas of the gasturbine engine, and a radial retention portion configured to engagerotor 10. Root 14A comprises a generally uniform cross sectional area inthe axial direction. As such, each segment of root 14A comprises agenerally equivalent root portion. When segments 26A, 26B and 26C areaxially aligned, root 14A can be axially inserted into slot 18 so thatdovetail engagement 22 (FIG. 1) is properly seated. Specifically, root12A includes channel 46 that axially receives ribs on slot 16 (FIG. 1)to prevent displacement of platform 14 in the radial direction. Root 14Aalso includes assembly bore 48 that is used to axially assemble andrestrain segments 26A-26C.

Stage 14B includes leading edge 50, trailing edge 52, convex side 54 andconcave side 56, and extends in a generally arcuate manner in the axialdirection. Stage 14B also extends generally radially from root 14A suchthat each of segments 26A, 26B and 26C include a small slice of thearcutate profile. Leading edge segment 26A and trailing edge segment26B, however, also include axially and tangentially extending portions58 and 60, respectively, that form forward and aft portions of the innerdiameter shroud that encircles blades 12. Leading edge 50 and trailingedge 52 of segments 26A and 26B, respectively, are planar such that theyare configured to be adjacent platforms from adjacent turbine stages.The radially outer surfaces 62 of segments 26A, 26B and 26C form agenerally flat flow path that is generally perpendicular to the hot gaspath. Convex side 54 has a curvature that is opposite that of pressureside 36 such that convex side 54 and pressure side 36 engage flushlywhen assembled with rotor 10. Likewise, concave side 56 has a curvaturethat is opposite that of suction side 38 such that concave side 56 andsuction side 38 engage flushly when assembled with rotor 10. Middlesegments 26C include planar forward and rear surfaces so that they canbe axially stacked or aligned between leading edge segment 26A andtrailing edge segment 26B to complete the shape of a platform.

Blade 12 and platform 14 are divided into elongate, thin segments tofacilitate processing of the CMC material from which they are produced.Specifically, segments 24A-24C of blade 12 and segments 26A-26C ofplatform 14 are comprised of laminar layers of ceramic fibers such thatthe fibers provide radial tensile strength to each segment and such thateach segment is properly densified with a liquid ceramic matrix. Blade12 and platform 14 are produced from similarly configured CMC plates,the specifics of which are described with reference to middle segment26C of platform 14 in FIGS. 5A and 5B. However, CMC plates of such aconstruction are suitable for use in any of segments 24A-24C andsegments 26A-26C.

FIG. 5A shows one of CMC blade platform segments 26C of FIG. 4fabricated from uni-directional ceramic fibers 64. FIG. 5B shows one ofCMC blade platform segments 26C of FIG. 4 fabricated from bi-directionalfibers 66 and 68. Segment 26C is comprised of a generally planar platehaving through-thickness t. Through-thickness t extends generally in theaxial direction when segment 26C is inserted into slot 18 (FIG. 1) ofrotor 10. Each plate is comprised of ceramic fiber laminar sheets thatspans root portion 14A and gas path portion 14B. In FIG. 5A, each sheet70 is comprised of a plurality of ceramic fiber strands 64 that areprovided as a uni-directional tape. In FIG. 5B, each sheet 72 iscomprised of a plurality of fiber strands 66 and 68 that are provided asa bi-directional woven mesh. Fibers 64, 66 and 68 comprise longcontinuous strands of lightweight and high strength materials such ascarbon or silicon carbide. Sheets 70 and 72 extend in a radial andtangential plane when segment 26C is inserted into slot 18 (FIG. 1) ofrotor 10. Sheets 70 and 72 are infused with a matrix of liquid ceramic,such as alumina or mullite. Segment 26C is baked or heated to solidifyand harden the liquid ceramic matrix into a lightweight and heatresistant ceramic solid having fibers 64-68 to provide reinforcement.Segment 26C is machined, such as with a diamond tip tool, to produceflat surfaces, curves for convex side 54 and concave side 56, and otherfeatures such as channel 46 and bore 48. Machining may take place afterthe liquid ceramic matrix is fully hardened or when the matrix issemi-hardened or in a green state. As such, platform 14 can be producedby individually shaping CMC plates and then assembling them to produceplatform 14, or by assembling individual CMC plates and then machiningthem as a unitary block to produce platform 14.

Configured as such, segment 26C includes material properties suitablefor use in rotary components of a turbine section of a gas turbineengine. Specifically, segment 26C has radial tensile strength capable ofwithstanding centrifugal forces generated by rotation of rotor 10 duringoperation of the gas turbine engine. For example, during operation of agas turbine engine, turbine blades are loaded with tension and bendingstresses that require maximum strength in the radial direction. Strands64 and 66 of sheets 70 and 72, respectively, extend continuously fromroot portion 14A through to gas path portion 14B in the radialdirection. As such, strands 64 and 66 are aligned in the direction inwhich segment 26C is subjected to the most stress. Furthermore, each ofsheets 70 and 72 is individually connected to rotor 10 at dovetail 22.Thus, the need for further radial strength enhancement, such as radialcompression means, is reduced or eliminated. Additionally, strands 64and 66 are free from sharp bends or folds, which reduces stresses fromaccumulating within the typically non-compliant fibers.

The segmented nature of platform 14, however, enables different segmentsto include folds that facilitate production of segments having angulargeometries. For example, leading edge segment 26A and trailing edgesegment 26B include fibers having bends to produce platform portions 58and 60 (FIGS. 3 and 4). However, platform portions 58 and 60 areproduced with bends that are approximately ninety degrees, which avoidssharp bends that tend to produce stress concentrations. Additionally,the masses of platform portions 58 and 60 are small such that thebending stresses at the fold are minimal. Thus, platform 14 is dividedinto thin, plate-like segments that have ceramic fiber strands that arealigned in a preferred orientation which results in optimal tensilestrength.

In various embodiments of the present invention, different componentsand segments can be made from CMC material or from traditional metalalloys to take advantage of different cost, weight and manufacturingbenefits of each material. For example, platforms comprise approximatelyone quarter of the weight of each blade, while metal alloys are easierto manufacture. Thus, in one embodiment, blade 12 is made from a highstrength alloy, while platform 14 is made from CMC material. Likewise,different segments of each component can be comprised of differentmaterials. For example, leading edge segments 24A and 26A and trailingedge segments 24B and 26B can be made from metal alloys to facilitatemanufacture of the more complex geometry of those parts, while middlesegments 24C and 26C can be made from CMC material to save weight. Thus,each component and each segment can be made from combinations of alloyand CMC material to optimize cost, weight, strength and performance ofeach segment.

The thin, plate-like segments also enable the liquid ceramic matrix tobe fully infused into the ceramic fibers comprising each segment. Liquidceramic infiltration typically takes place under high pressures toenable the infused liquid to pass through multiple layers of ceramicfibers. Each ceramic fiber layer is approximately 10 to approximately 20mils (˜0.025-˜0.05 cm) thick, thus requiring tens or hundreds of layersto produce parts having adequate thickness and strength. However, thicklayering of the ceramic fibers, particularly at areas where the geometrychanges shape such as at sharp bends that requires the ceramic matrix tochange course during the infiltration, prevents the matrix fromuniformly penetrating the fibers. Through-thickness t of segment 26C isselected such that the liquid ceramic matrix can penetrate the ceramicfibers to achieve a uniform density and porosity. Porosity isundesirable as it produces discontinuities in the solidified matrix thatinhibits heat distribution throughout the CMC material and that providesan initiation point for crack propagation. In one embodiment of theinvention, each segment has a maximum thickness of approximately 0.25inches (˜0.64 cm). Different segments may have different thicknesses toprovide more or less radial strength at different positions along theaxial length of platform 14 or to obtain a desired overall axial lengthof platform 14. Flat-plate construction also has other benefits as isknown in the art, such as reduction in anisotropic shrinkage, reductionof delamination flaws, and dimensional control.

FIG. 6 shows a partially cut away view of an embodiment of a retentionmechanism for use with removable CMC blade platform 14 of FIGS. 3 and 4.Blade 12 and platform 14 are seated within slots 16 and 18,respectively, such that pressure side 36 (FIG. 2) of blade 12 flushlyabuts convex side 54 (FIG. 3) of platform 14. Forward retention plates74A and 74B and aft retention plate 74C axially restrain blade 12 andplatform 14 within slots 16 and 18. Segments 26A-26C are preassembled inan axial stack using a glue or adhesive such that platform 14 can beinserted into slot 18 of rotor 10. The glue temporarily holds segments26A-26C together to facilitate assembly of cover plates 74A, 74B and 74Cwith rotor 12 using fastener 76. The glue ultimately burns off duringoperation of the gas turbine engine.

Each retention plate comprises a segment of an annular ring that, whenassembled, has an outer diameter that is positioned just below the outerextents of slots 16 and 18, and an inner diameter that is positionedjust below the troughs of slots 16. Retention plates 74A, 74B and 74Care secured to rotor 10 using fastener 76, which is extended throughbores within plates 74B and 74C and bore 48 within platform root 14A.Fastener 76 directly compresses segments 26A-26C between plates 74B and74C to keep segments 26A-26C axially assembled to each other, and tokeep root 14A axially restrained within slot 18. Retention plates 74Band 74C also extend tangentially away from slot 18 to cover portions ofadjacent slots 16. Retention plates 74B and 74C therefore also providecompressive forces to blade root 12A to maintain assembly and to provideaxial restraint. Secured as such, the retention plates prevent roots 12Aand 14A from moving axially within slots 16 and 18.

The retention plates also control air flow through slots 16 and 18 ofrotor 10. For example, retention plates 74A and 74B include cut-outs 78Aand 78B, which form an opening to permit cooling air into slot 16. Thus,the retention plates prevent cooling air from passing through rotor 10between root 12A and slot 16 and root 14A and slot 18, but permitcooling air to pass underneath root 12A of blade 12 so as to be able toenter any cooling passages extending through airfoil 12B and root 12A,such as cavity 44 (FIG. 2).

FIG. 7 shows another embodiment of a retention mechanism for use withremovable CMC blade platform 14 of the present invention. In theembodiment of FIG. 7, slot 18 of FIG. 6 is replaced with slot 80 andradial flanges or tabs 82A and 82B. Slot 80 comprises a tangential slotbetween tabs 82A and 82B and extends clear through to adjacent slots 16.Slot 80 and middle segments 26C are not provided with a radial retentionengagement, such as a dovetail, but are rather radially restrained byfastener 84. Fastener 84 extends through leading edge segment 26A, tab82A, middle segments 26C, tab 82B and trailing edge segment 26B tomaintain axial assembly of segments 26A-26C and to radially restrainplatform 14. Because of tabs 82A and 82B, fastener 84 does not providedirect compression to middle segments 26C. Leading edge segment 26A andtrailing edge segment 26B are, however, provided with axial tabs 86A and86B, respectively, which compress middle segments 26C when fastener 84is put in tension. Axial tabs 86A and 86B extend axially from gas pathportions 14B of segments 26A and 26B to span the gaps produced by radialtabs 82A and 82B between segments 26A and 26B and middle segments 26C,respectively. Segments 26A and 26B also include tabs or wings 88A and88B, which extend tangentially from root portions 14A to provide axialrestraint to adjacent blade roots 12A. For example, leading edge segment26A includes wing 88A that provides forward axial restraint to blade 12,and trailing edge segment 26B includes wing 88B that provides rearwardaxial restraint to a tangentially opposite adjacent blade. However, inother embodiments, segments 26A and 26B may include any combination offorward, rearward and tangentially opposing blade retention wings. Blade12 is individually radially restrained by a radial retention engagementsuch as fir tree engagement 20. Assembled as such, leading edge segment26A, trailing edge segment 26B and fastener 84 provide axial retentionto both blade 12 and platform 14, and a continuous ring of axialretention plates are not needed.

The CMC segments of the present invention allow for directionalalignment of reinforcing fibers so that material strength properties canbe matched to performance needs. Furthermore, the matrix material isuniformly infused into the reinforcing fibers to achieve maximum heattransfer properties as well as to avoid weaknesses associated withporosity. Thus, CMC segmented components are highly suitable for use asrotary components in a gas turbine engine, as is described with respectto FIGS. 1-7, where light weight, heat resistance and high radialstrength is desirable. Additionally, CMC segmented components are highlysuitable for use in stationary components within gas turbine engines.

FIG. 8 shows an embodiment of a stationary segmented CMC vane.Stationary CMC segmented vane 90 comprises airfoil 92 and split shrouds94A, 94B and 94C. Vane 90 and shrouds 94A, 94B and 94C are produced asseparate pieces such that each piece can be produced from materials andwith methods that result in optimal performance. In one embodiment,airfoil 92 is formed from CMC material segments to enhance strength,heat and weight capabilities of airfoil 92, and shrouds 94A-94C are castfrom high strength alloys for ease of manufacture. A plurality of vanes90 and shrouds are assembled to form an array of stationary vanesdisposed circumferentially about a centerline of a gas turbine engine.For example, each vane 90 would be positioned adjacent a pair of innerdiameter shrouds 94A and 94B, and a pair of outer diameter shrouds 94C.Thus, for each vane in the array, there is a corresponding inner andouter diameter shroud.

Airfoil 92 is divided into CMC segments including leading edge segment96A, trailing edge segment 96B and a plurality of middle segments 96C.In the embodiment shown, airfoil 92 includes eight middle segments 96C.As is described with respect to segment 26C of FIG. 5A, segments 96A-96Care formed of thin layers of reinforcing fibers having a thickness suchthat the fibers can be completely infused with a ceramic matrixcomposite material. Additionally, the fibers are radially oriented suchthat segments 96A-96C have high tensile strength. Airfoil 92 and shrouds94A, 94B and 94C are shaped to allow for thermal growth between segments96A, 96B and 96C and shrouds 94A, 94B and 94C from heat generated duringoperation of the gas turbine engine. In other embodiments, leading edgesegment 96A and trailing edge segment 96B are comprised of a metal alloymaterial.

Airfoil 92 is configured to impinge hot gas in a gas turbine engine andincludes leading edge 98, trailing edge 100, concave pressure side 102and convex suction side 104. Airfoil 92 extends in a generally arcuate,or airfoil shape, manner in the axial direction. Airfoil 92 also extendsgenerally radially from inner diameter shrouds 94A and 94B to outerdiameter shroud 94C such that each of segments 96A, 96B and 96C includesa small slice of the profile of airfoil 92. Leading edge segment 96Aincludes a rounded forward surface that forms leading edge 98 and isdesigned to impinge an oncoming gas stream. Leading edge segment 96Aalso includes a planar rear surface designed to abut one of middlesegments 96C. Trailing edge segment 96B includes a rounded rear surfacethat forms trailing edge 100 and is designed to minimize flow separationof the gas stream. Trailing edge segment 96B also includes a planarforward surface designed to abut one of middle segments 96C. Middlesegments 96C include planar forward and rear surfaces so that they canbe axially stacked or aligned between leading edge segment 96A andtrailing edge segment 96B to complete the shape of an airfoil. The sidesof middle segments 96C include small segments of pressure side 102 andsuction side 104 to complete the airfoil shape. Additionally, eachsegment can be split into pressure side and suction side segment halvesto facilitate assembly with shrouds 94A-94C, and to produce hollowairfoils to minimize weight, as is done with blade 12 of FIG. 2. Airfoil92 also includes groove 106 such that segments 96A-96C can be assembledwith shrouds 94A-94C. Groove 106 can be machined into segments 96A-96Cor can be cast into the CMC segments.

Shrouds 94A-94C are comprised of metal alloys to facilitatemanufacturing of the complex shapes of shrouds 94A-94C. Shrouds 94A-94Care shaped to seal between adjacent airfoils 92. The shrouds prevent hotgas that flows over airfoil 92 from leaking into the gas turbine engineand maintain hot gases flowing axially through stages of vanes andblades in the gas turbine engine. For example, shroud 94C includesleading edge side 108, trailing edge side 110, pressure side 112 andsuction side 114. Leading edge side 108 and trailing edge side 110include flanges that can be mounted to a casing structure in a gasturbine engine. For example, leading edge side 108 includes hole 116through which a threaded fastener can be extended to secure shroud 94Cto a casing structure. Pressure side 112 and suction side 114 havecontours that match pressure side 102 and suction side 104 of airfoil92, respectively. Additionally, leading edge side 108 and trailing edgeside 110 abut at leading edge 98 and trailing edge 100, respectively,such that a continuous hoop-shaped shroud structure is formed. Pressureside 112 and suction side 114 each also include a lip from which airfoil92 is suspended. For example, suction side 112 includes lip 118 that isinserted into groove 106 along suction side 104 of airfoil 92. Outerdiameter shrouds 94C are connected together, such as with fasteners, toform a rigid structure that supports the outer diameter ends of airfoils92. As outer diameter shrouds 94C are connected to a casing structure,segments 96A, 96B and 96C are clamped between outer diameter shroudsegments.

Segments 96A-96C are preassembled in an axial stack using a glue oradhesive such that airfoil 92 can be connected with outer and innerdiameter shrouds. The glue temporarily holds segments 96A-96C togetherto facilitate assembly of the shrouds. The glue ultimately burns offduring operation of the gas turbine engine. In other embodiments, athreaded fastener is positioned within a bore extending through segments96A-96C either radially inward of inner diameter shrouds or radiallyoutward of outer diameter shrouds. Inner diameter shroud segments, suchas shrouds 94A and 94B, are connected to inner diameter ends of airfoils92 using a similar lip and groove system as is used with outer diametershrouds 94C. The inner diameter shroud segments are connected to eachother using, for example, threaded fasteners to provide a rigidstructure that supports the inner diameter ends of airfoils 92. Theassembled inner diameter shroud is then supported by another supportstructure, such as by a sealing system or a cooling system supported bya bearing surrounding the shaft of the gas turbine engine.

FIG. 9 shows another embodiment of a stationary segmented CMC vane.Stationary CMC segmented vane 120 comprises airfoil 122, split shrouds124A and 124B, and retention tabs 126A and 126B. Stationary vane 120functions similar to that of stationary vane 90 of FIG. 8 in thatshrouds 124A and 124B support airfoil 122 within a gas turbine engine ina stationary manner. Airfoil 122 and shrouds 124A and 124B arefabricated as separate pieces such that each can be made to performoptimally. For example, shrouds 124A and 124B are fabricated from alloymaterials for ease of manufacture and for strength. Airfoil 122 iscomprised of CMC segments 128A, 128B and 128C. Segments 128A-128C areshaped and formed using similar methods as are used to produce segments96A-96C of FIG. 8 and segment 26C of FIG. 5A. The outer and innerdiameter ends of airfoil 122 are, however, shaped differently tointeract with shrouds 124A and 124B.

Airfoil 122 is shaped to impinge hot gas within a gas turbine engine andincludes leading edge 130, trailing edge 132, pressure side 134 and asuction side (not visible in FIG. 9), that form an airfoil shape.Segments 128A-128C include planar forward and rear surfaces that permitsegments 128A-128C to be stacked axially. Leading edge segment 128Aincludes leading edge 130, and trailing edge segment 128B includestrailing edge 132, which include rounded surfaces for aerodynamic flowpurposes. Middle segments 128C include curved surfaces that formpressure side 134 and the suction side when stacked. Outer diameter end136 includes retention lug 138 that interacts with retention tabs 126Aand 126B to secure airfoil 122 to outer diameter shroud 124A. Similarly,the inner diameter end of airfoil 122 interacts with a pair of retentiontabs radially inward of inner diameter shroud 124B.

Shrouds 124A and 124B perform similar functions to that of shrouds94A-94C of FIG. 8, but, rather than filling gaps between adjacentairfoils, surround an end of an airfoil such that each shroud contactsonly one airfoil. Specifically, shroud 124A engages only outer diameterend 136 of one airfoil 122, but contacts two identical shrouds atpressure side wall 140 and suction side wall 142. Side walls 140 and 142are generally planar such that the walls abut flush with adjacentshrouds. Side walls 140 and 142 also include slots for receiving featherseals. For example, pressure side 140 includes slot 144 that receives athin sheet metal-like piece that extends into a slot on an adjacentshroud to inhibit air from escaping the gas path along airfoils 122within a gas turbine engine. Shrouds 124A and 124B are conventionallymounted within the gas turbine engine, such as by using bore 146 andthreaded fasteners. As such, shrouds 124A and 124B connect with othershrouds to form rigid hoop-like structures for supporting airfoils 122.Shrouds 124A and 124B, however, include axially extending openings forengaging retention lug, such as retention lug 138, of airfoil 122.

FIG. 10 shows airfoil 122, outer diameter vane shroud 124A and retentiontabs 126A and 126B of vane 120. Vane shroud 124A and retention tab 126Aare broken away to show the interaction of airfoil 122 with retentiontabs 126A and 126B. Additionally, leading edge segment 128A and aportion of middle segments 128C are omitted from FIG. 10. Outer diameterend 136 of airfoil 122 includes retention lug 138, which includespressure side groove 146A and suction side groove 146B. Shroud 124Aincludes pressure side wall 140, suction side wall 142, seal slot 144,trailing edge wall 148, pressure side floor 150, suction side floor 152and opening 154.

Opening 154 extends generally diagonally across shroud 124A to separatepressure side floor 150 from suction side floor 152. Opening 154 extendsacross the length of shroud 124A between trailing edge wall 148 and acorresponding leading edge wall (not shown). Thus, floors 150 and 152comprise generally triangular shaped segments that are bordered by walls140, 142 and 148 and opening 154. Retention lug 138 comprises agenerally rectangular projection that extends from the distal, orradially outer, surfaces of the portions of CMC segments 128A-128Cforming airfoil 122. Retention lug 138 is integrally cast in CMCsegments 128A-128C to preserve the radial integrity and strength ofairfoil 122. Lug 138 has the same general shape and size, or crosssection, of opening 154 such that lug 138 fits tightly into opening 154.Pressure side groove 146A and suction side groove 146B are positionedalong lug 138 at a specified distance above the distal surfaces of CMCsegments 128A-128C that form airfoil 122. The distance generallycorresponds to a thickness of floors 150 and 152. Grooves 146A and 146Bcomprise generally C-shaped channels that have flat radially innersurfaces and outer surfaces that define a channel height. The channelheight generally corresponds to the thickness of retention tabs 126A and126B.

When lug 138 is inserted into opening 154, floors 150 and 152 will reston the radial outer surfaces of segments 128A-128C comprising airfoil122. Floors 150 and 152 include notches such that shroud 124A can befitted onto segments 128A-128C while lug 138 is fitted into opening 154.For example, floor 150 includes notch 156 that traces the contour ofpressure side 134 of airfoil 122. The radially inner surfaces of grooves146A and 146B are generally coplanar with the radially outer surfaces offloors 150 and 152, respectively. The radially outer surfaces of grooves146A and 146B are generally coplanar with the radially outer surfaces ofretention tabs 126A and 126B, when retention tabs 126A and 126B rest onfloors 150 and 152. With retention tabs 126A and 126B resting on floors150 and 152, retention tabs 126A and 126B can be inserted into grooves146A and 146B, respectively. Retention tabs 126A and 126B havethicknesses that fill grooves 146A and 146B such that tabs 126A and 126Bare tightly fitted. Retention tabs 126A and 126B have widths that allowtabs 126A and 126B to extend completely into grooves 146A and 146B aswell as to overhang floors 150 and 152. Retention tabs 126A and 126Bhave lengths that extend from trailing edge wall 148 to thecorresponding leading edge wall. As such, retention tabs 126A and 126Bhave a general parallelogram shape that permits each tab to fit betweenlug 138, side walls 140 and 142, trailing edge wall 148 and thecorresponding leading edge wall. With retention tabs 154A and 154Bseated in grooves 146A and 146B and resting on floors 150 and 152,respectively, airfoil 122 is restrained from moving either radiallyinward or radially outward with respect to shroud 124A. Retention tabs126A and 126B are welded in place, or otherwise secured, to preventdislodgement. Inner diameter shroud 124B (FIG. 9) is assembled toairfoil 122 in a similar fashion. Shrouds 124A and 124B are connected toairfoil 122 such that vanes 120 can be subsequently installed into a gasturbine engine.

Thus, thin segments of CMC material can be shaped and stacked to formlarger components having more complex shapes. The thin segments of CMCmaterial allow the ceramic matrix material to be properly infused intoreinforcing fibers to gain the full benefit of the ceramic matrix. Thethin segments also allow the reinforcing fibers to be aligned in adirection that will provide maximum strength to the assembled component.The components can then be used in high heat, high stress environmentswhere weight limitations are of concern. For example, CMC segments ofthe present invention can be shaped and assembled to form variouscomponents for a hot section of a gas turbine engine, such as blades,vanes, platforms and shrouds.

Although the present invention has been described with reference topreferred embodiments, workers skilled in the art will recognize thatchanges may be made in form and detail without departing from the spiritand scope of the invention.

1. A segmented component for a gas turbine engine, the componentcomprising: a gas path portion for interacting with gas flow in the gasturbine engine, the gas path portion comprising: a forward portionforming a leading edge of a stationary vane; an aft portion forming atrailing edge of the stationary vane; and a plurality of middle portionsforming a pressure side and a suction side of the stationary vane;wherein the component is divided into axially aligned segmentscomprising: a forward segment; an aft segment; and a plurality of middlesegments disposed between the forward segment and the aft segment, theplurality of middle segments comprising radially elongate ceramic matrixcomposite material plates.
 2. The segmented component of claim 1 whereineach of the segments are comprised of ceramic fiber and have a thicknesssuch that an entirety of the fiber can be nearly uniformly infiltratedwith a ceramic matrix liquid.
 3. The segmented component of claim 2wherein each of the plurality of middle segments are comprised ofceramic matrix composite plates having an axial thickness up toapproximately 0.25 inches (˜0.635 cm).
 4. The segmented component ofclaim 1 wherein each of the segments is comprised of a ceramic matrixcomposite plate having fabric mesh fibers extending in a radial andtangential plane with respect to a centerline of the gas turbine engine.5. The segmented component of claim 1 and further comprising: anon-integral outer diameter shroud disposed at an outer diameter end ofthe stationary vane; and a non-integral inner diameter shroud disposedat an inner diameter end of the stationary vane.
 6. The segmentedcomponent of claim 5 wherein: the outer diameter shroud and the innerdiameter shroud include a contoured surface configured to match acontour of either the pressure side or the suction side of thestationary vane, the contoured surfaces having a lip; and the innerdiameter end and the outer diameter end of the stationary vane eachinclude a groove into which the lips on the contoured surface fit,respectively.
 7. The segmented component of claim 5 and furthercomprising: first and second retention tabs disposed against the outerdiameter shroud and inner diameter shroud, respectively; wherein: theouter diameter shroud and the inner diameter shroud each include anopening extending radially through the shroud; the inner diameter endand the outer diameter end of the stationary vane each include a flangeextending into an opening, the flanges each having a channel; and thefirst and second retention tabs are disposed within a channel.
 8. Thesegmented component of claim 1 wherein the gas path portion furthercomprises: a first diametric end; a second diametric end; and aretention lug disposed at the first diametric end.
 9. The segmentedcomponent of claim 8 and further comprising a groove extending acrossthe retention lug.
 10. The segmented component of claim 8 and furthercomprising: a shroud positioned at the first diametric end of the gaspath portion; and a retention tab coupled to the retention lug andjoined to the shroud.
 11. The segmented component of claim 10 whereinthe shroud is comprised of an alloy material.
 12. The segmentedcomponent of claim 11 wherein the retention tab is welded to the shroud.13. The segmented component of claim 12 wherein the shroud comprises: anopening through which the retention lug extends.
 14. The segmentedcomponent of claim 13 wherein the shroud further comprises: a notchdisposed along the opening that traces a contour of the gas pathportion.
 15. The segmented component of claim 1 wherein the gas pathportion further comprises: a first diametric end; a second diametricend; and a groove circumscribing the gas path portion proximate thefirst diametric end.
 16. The segmented component of claim 15 and furthercomprising: a shroud coupled to the first diametric end of the gas pathportion, the shroud comprising: an opening through which the firstdiametric end of the gas path portion extends; and a lip extending fromthe opening to engage the groove.
 17. A segmented ceramic matrixcomposite turbine vane for a gas turbine engine, the turbine vanecomprising: a gas path portion for interacting with gas flow from thegas turbine engine, the gas path portion being divided into axiallyaligned segments comprising: a forward segment forming a leading edge ofa stationary vane; an aft segment forming a trailing edge of thestationary vane; and a plurality of middle segments forming a pressureside and a suction side of the stationary vane; wherein the plurality ofmiddle segments comprise radially elongate ceramic matrix compositematerial plates having fabric mesh fibers extending in a radial andtangential plane with respect to a centerline of the gas turbine engine.18. The segmented ceramic matrix composite turbine vane of claim 17 andfurther comprising: a retention lug disposed at a first diametric end ofthe gas path portion; and a groove extending across the retention lug.19. The segmented ceramic matrix composite turbine vane of claim 17 andfurther comprising a groove circumscribing the gas path portionproximate a first diametric end of the gas path portion.
 20. A segmentedcomponent for a gas turbine engine, the component comprising: a gas pathportion for extending into a gas flow within a gas turbine engine, thegas path portion comprising a leading edge, a trailing edge, a pressureside and a suction side of an airfoil; wherein the gas path portion isdivided into axially aligned segments comprising: a forward segment; anaft segment; and a plurality of middle segments disposed between theforward segment and the aft segment, each of the plurality of middlesegments comprising a radially elongate ceramic matrix compositematerial plate having reinforcing fibers extending radially through thegas path portion, and stacked axially to a thickness that permitsuniform infiltration of a ceramic matrix liquid.